Typical gas turbine engine fuel supply systems include a fuel source, such as a fuel tank, and a main fuel pump that receives fuel drawn from the fuel source and delivers pressurized fuel to the fuel manifolds in the engine combustor via a fuel supply line. The main fuel pump is typically implemented using a positive displacement pump that is driven directly by the engine gearbox. Thus, the fuel flow supplied by the main fuel pump is proportional to engine speed.
The fixed displacement of a main fuel pump is typically sized to produce the fuel flow that is needed to run the engine at a maximum demand case, which is typically during engine start-up, where engine speed is relatively low, or during takeoff, where fuel demand is relatively high. As such, at other operating conditions, such as idle or high altitude cruise, the main fuel pump supplies much more fuel than the engine needs.
The known fuel supply systems described above generally operate safely and robustly, but can exhibit certain drawbacks. For example, the overcapacity of the main fuel pump results in increased horsepower extraction, which increases engine fuel consumption. This overcapacity also increases the overall fuel temperature within the fuel supply system.
Hence, there is a need for a gas turbine engine fuel supply system that reduces fuel pumping overcapacity and leads to a decrease in overall engine fuel consumption and overall fuel temperature. The present invention addresses at least these needs.